Composite honeycomb sandwich panel splicing

ABSTRACT

A method and apparatus for joining composite parts is present. A first composite part is aligned with a second composite part to form an aligned structure with a splice line. An adhesive layer may cover the splice line. A composite material is placed on the adhesive layer. The aligned structure is cured with the adhesive layer and the composite material to form a joined structure.

BACKGROUND INFORMATION

1. Field

The present disclosure relates generally to manufacturing and, inparticular, to a method and apparatus for joining parts to each other.Still more particularly, the present disclosure relates to a method andapparatus for joining composite parts to each other.

2. Background

Aircraft are being designed and manufactured with greater and greaterpercentages of composite materials. Some aircraft may have more thanfifty percent of their primary structure made from composite materials.Composite materials may be used in aircraft to decrease the weight ofthe aircraft. This decreased weight may improve payload capacities andfuel efficiencies. Further, composite materials may provide longerservice life for various components in an aircraft.

Composite materials may be strong, light-weight materials, created bycombining two or more dissimilar components. For example, withoutlimitation, a composite may include fibers and resins. The fibers andresins may be combined to form a cured composite material.

Composite materials such as, for example, without limitation, compositehoneycomb sandwich panels may be used in the interior of an aircraft.These panels may be used for walls, floors, light covers, and othersuitable structures for the interior of an aircraft. A compositesandwich panel may have a honeycomb core with a faceplate (face sheet)on either side of the honeycomb core. This faceplate may be, forexample, without limitation, a laminate and/or other composite material.

In some cases, it may be desirable to join or attach composite parts toeach other. For example, without limitation, with light shields, thesetypes of panels may be made of a carbon fiber skin with a honeycombcore. The ends of these light shields may be rounded. These panels alsomay have different length requirements depending on the interiorconfiguration of a passenger cabin.

Rather than purchasing and/or manufacturing a tool for each length, atool with the longest length may be used to create a panel. This panelmay then be cut into parts of the appropriate size. These parts may bejoined to each other with a structural bond. This joining of parts mayalso be referred to as splicing. This process may involve the use ofnon-robust materials and may be tedious. As a result, increased out oftolerance and scrap rates may occur.

The current process may involve removing a portion of the honeycomb coreout of each part to be joined to each other. The two parts may then bejoined by applying a potting and/or adhesive material within the core.The parts may be clamped to each other for several hours. Thereafter,the parts may be sanded for the application of a decorative layer.

For example, the parts may be cured for around 24 hours to fully jointhe parts to each other. This type of process may have less thandesirable out of tolerance rates during assembly and/or installation ofthese panels. The time needed to join the parts to each other also mayslow down the speed at which an aircraft may be assembled as well aspossibly increase assembly costs.

Therefore, it would be advantageous to have a method and apparatus thatovercomes one or more of the issues discussed above.

SUMMARY

In one advantageous embodiment, a method for joining composite parts maybe present. A first composite part may be aligned with a secondcomposite part to form an aligned structure with a splice line. Anadhesive layer may cover the splice line. A composite material may beplaced on the adhesive layer. The aligned structure may be cured withthe adhesive layer and the composite material to form a joinedstructure.

In another advantageous embodiment, a method for manufacturing acomposite honeycomb sandwich panel may be present. A first compositehoneycomb sandwich panel may be aligned with a second compositehoneycomb sandwich panel to form an aligned structure with a spliceline. An epoxy film may be placed on the splice line. A fabric with anembedded resin may be placed on the epoxy film. The aligned structuremay have a first side and a second side. The epoxy film and the fabricwith the embedded resin may be placed on the first side and the secondside. The fabric may be selected from one of a fiberglass fabric and acarbon fabric. The resin may be selected from one of a phenolic resinand an epoxy resin. The aligned structure with the epoxy film and thefabric with the embedded resin may be placed in a die. The die may coverthe splice line. The aligned structure in the die may be heated underpressure to form a joined structure. A decorative layer may be appliedon the joined structure. The completed structure may be selected fromone of a composite panel, a wall panel, and a light shield.

In yet another advantageous embodiment, an apparatus may comprise afirst composite part, a second composite part, an adhesive layer, acomposite material, and a tool. The first composite part may be alignedwith a second composite part to form an aligned structure with a spliceline. The adhesive layer may cover the splice line. The compositematerial may be located on the adhesive layer. The tool may be capableof holding the aligned structure with the adhesive layer covering thesplice line and the composite material located on the adhesive layer.

In still yet another advantageous embodiment, an apparatus for splicinghoneycomb composite parts may comprise a first composite honeycombsandwich panel, a second composite honeycomb sandwich panel, an epoxyfilm, a fabric, and a tool. The first composite honeycomb sandwich panelmay be aligned with a second composite honeycomb sandwich panel to forman aligned structure with a splice line. The epoxy film may cover thesplice line. The fabric with an embedded resin may be located on theepoxy film. The tool may be capable of holding the aligned structurewith the epoxy film located on the splice line and the fabric with theembedded resin. The fabric may be selected from one of a fiberglassfabric and a carbon fabric. The resin may be selected from one of aphenolic resin and an epoxy resin. The joined structure may be selectedfrom one of a composite panel, a wall panel, and a light shield.

The features, functions, and advantages may be achieved independently invarious embodiments of the present disclosure or may be combined in yetother embodiments in which further details can be seen with reference tothe following description and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The novel features believed characteristic of the advantageousembodiments are set forth in the appended claims. The advantageousembodiments, however, as well as a preferred mode of use, furtherobjectives, and advantages thereof, will best be understood by referenceto the following detailed description of an advantageous embodiment ofthe present disclosure when read in conjunction with the accompanyingdrawings, wherein:

FIG. 1 is a diagram illustrating an aircraft manufacturing and servicemethod in accordance with an advantageous embodiment;

FIG. 2 is a diagram of an aircraft in which an advantageous embodimentmay be implemented;

FIG. 3 is a diagram of a splicing environment in accordance with anadvantageous embodiment;

FIG. 4 is a diagram of a cross-sectional view of an aligned structure inaccordance with an advantageous embodiment;

FIG. 5 is a diagram of a cross-sectional view illustrating layers ofmaterials for joining composite parts in accordance with an advantageousembodiment;

FIG. 6 is a diagram of a cross-sectional view illustrating a completedstructure in accordance with an advantageous embodiment;

FIG. 7 is a diagram of a composite honeycomb panel in accordance with anadvantageous embodiment;

FIG. 8 is a diagram illustrating application of an adhesive layer inaccordance with an advantageous embodiment;

FIG. 9 is a diagram illustrating an application of a composite materialto a structure in accordance with an advantageous embodiment;

FIG. 10 is a diagram illustrating application of a composite material toa structure in accordance with an advantageous embodiment;

FIG. 11 is a diagram illustrating an aligned structure on a tool inaccordance with an advantageous embodiment;

FIG. 12 is a diagram illustrating a cured aligned structure inaccordance with an advantageous embodiment;

FIG. 13 is a diagram of an aligned structure in accordance with anadvantageous embodiment; and

FIG. 14 is a flowchart of a process for joining composite parts inaccordance with an advantageous embodiment.

DETAILED DESCRIPTION

Referring more particularly to the drawings, embodiments of thedisclosure may be described in the context of aircraft manufacturing andservice method 100 as shown in FIG. 1 and aircraft 200 as shown in FIG.2. Turning first to FIG. 1, a diagram illustrating an aircraftmanufacturing and service method is depicted in accordance with anadvantageous embodiment. During pre-production, exemplary aircraftmanufacturing and service method 100 may include specification anddesign 102 of aircraft 200 in FIG. 2 and material procurement 104.

During production, component and subassembly manufacturing 106 andsystem integration 108 of aircraft 200 in FIG. 2 takes place.Thereafter, aircraft 200 in FIG. 2 may go through certification anddelivery 110 in order to be placed in service 112. While in service by acustomer, aircraft 200 in FIG. 2 is scheduled for routine maintenanceand service 114, which may include modification, reconfiguration,refurbishment, and other maintenance or service.

Each of the processes of aircraft manufacturing and service method 100may be performed or carried out by a system integrator, a third party,and/or an operator. In these examples, the operator may be a customer.For the purposes of this description, a system integrator may include,without limitation, any number of aircraft manufacturers andmajor-system subcontractors; a third party may include, withoutlimitation, any number of venders, subcontractors, and suppliers; and anoperator may be an airline, leasing company, military entity, serviceorganization, and so on.

With reference now to FIG. 2, a diagram of an aircraft is depicted inwhich an advantageous embodiment may be implemented. In this example,aircraft 200 is produced by aircraft manufacturing and service method100 in FIG. 1 and may include airframe 202 with a plurality of systems204 and interior 206. Examples of systems 204 include one or more ofpropulsion system 208, electrical system 210, hydraulic system 212, andenvironmental system 214. Any number of other systems may be included.Although an aerospace example is shown, different advantageousembodiments may be applied to other industries, such as the automotiveindustry.

Apparatus and methods embodied herein may be employed during any one ormore of the stages of aircraft manufacturing and service method 100 inFIG. 1. For example, components or subassemblies produced in componentand subassembly manufacturing 106 in FIG. 1 may be fabricated ormanufactured in a manner similar to components or subassemblies producedwhile aircraft 200 is in service 112 in FIG. 1.

Also, one or more apparatus embodiments, method embodiments, or acombination thereof may be utilized during production stages, such ascomponent and subassembly manufacturing 106 and system integration 108in FIG. 1, for example, without limitation, by substantially expeditingthe assembly of or reducing the cost of aircraft 200. Similarly, one ormore of apparatus embodiments, method embodiments, or a combinationthereof may be utilized while aircraft 200 is in service 112 or duringmaintenance and service 114 in FIG. 1.

For example, without limitation, advantageous embodiments may be usedfor drawing composite parts together during component and subassemblymanufacturing, system integration, and/or maintenance and service 114.The different advantageous embodiments may be used to join thesecomposite parts for use within interior 206 of aircraft 200 in theseillustrative examples.

The different advantageous embodiments recognize and take into accountthat the currently used processes for joining composite parts to eachother may be time consuming and costly. The different advantageousembodiments recognize and take into account that the currently usedprocesses may also result in higher than desired out of tolerance jointsof the joined parts.

Thus, the different advantageous embodiments provide a method andapparatus for joining composite parts to each other. A first compositepart may be aligned with a second composite part to form an alignedstructure with a splice line. An adhesive layer may be placed on thesplice line to cover the splice line. A composite material may be placedon the adhesive layer. The aligned structure may be cured with theadhesive layer and the composite material to form a joined structure.

With reference now to FIG. 3, a diagram of a splicing environment isdepicted in accordance with an advantageous embodiment. Splicingenvironment 300 may be used to join composite part 302 to composite part304. Composite part 302 may be aligned with composite part 304 to formaligned structure 306 with splice line 308. Aligned structure 306 may bean aircraft structure for use within interior 206 of aircraft 200 inthese illustrative examples.

In this depicted example, adhesive layer 310 may be placed onto surface312 of aligned structure 306 such that adhesive layer 310 covers spliceline 308. Adhesive layer 310 may be, for example, without limitation,adhesive film 314. Adhesive film 314 may be, for example, withoutlimitation, an epoxy film or some other suitable type of adhesive film.

Composite material 316 may be placed onto adhesive film 314. Compositematerial 316 may be fabric 318 impregnated with resin 320. In thedifferent advantageous embodiments, fabric 318 may be, for example,without limitation, a fiberglass fabric, a carbon fabric, and/or someother suitable fabric. Resin 320 may be, for example, withoutlimitation, a phenolic resin, an epoxy resin, or some other suitableresin.

In the different advantageous embodiments, tool 322 may secure alignedstructure 306, adhesive layer 310, and composite material 316 in place.Tool 322 may be, for example, without limitation, a die that may coveror secure only a portion of aligned structure 306 on which adhesivelayer 310 and composite material 316 may be located. Tool 322 may be,for example, without limitation, a contoured aluminum die, a caul plate,or some other suitable tool. Tool 322 may receive and/or hold alignedstructure 306 in place during a curing process. Tool 322 also mayprovide pressure by clamping aligned structure 306 before and/or duringthe curing process.

Aligned structure 306 with adhesive layer 310 and composite material 316may be placed on tool 322 and cured using heat source 324. Heat source324 may be implemented using a number of different types of devices. Forexample, without limitation, heat source 324 may be heating elementsattached to and/or embedded in tool 322. In other advantageousembodiments, tool 322 may be placed into heat source 324 when heatsource 324 takes the form of an autoclave oven. Of course, heat source324 may take other forms and may heat tool 322 and aligned structure 306in a number of different ways including, for example, withoutlimitation, heated oil, steam, electricity, and other suitable heatingmechanisms. Additionally, heat source 324 also may apply pressure duringthe curing process.

With the use of adhesive layer 310 and composite material 316, curingaligned structure 306 with adhesive layer 310 and composite material 316may only require around eight minutes instead of the currently used 24hours with current processes to splice parts to each other.

After aligned structure 306 has been cured with adhesive layer 310 andcomposite material 316, aligned structure 306 may become joinedstructure 326. Decorative layer 328 may be bonded onto joined structure326 to form completed structure 330.

The illustration of splicing environment 300 in FIG. 3 is not meant toimply architectural or physical limitations to the manner in whichdifferent advantageous embodiments may be implemented. Some advantageousembodiments may have other components in addition to, or in place of,the ones illustrated. In other advantageous embodiments, some componentsmay be unnecessary. For example, without limitation, additionalcomposite parts, in addition to composite parts 302 and 304, may bespliced together to form joined structure 326. In these examples,composite part 302 and composite part 304 may be similar structures suchas, for example, without limitation, light shields. In yet otheradvantageous embodiments, composite part 302 and composite part 304 maybe dissimilar structures joined to each other using adhesive layer 310and composite material 316.

In another illustrative example, splicing environment 300 may employanother heat source other than heat source 324. For example, a heatblanket and/or inductive heating also may be used, depending on theparticular implementation.

With reference now to FIG. 4, a diagram of a cross-sectional view of analigned structure is depicted in accordance with an advantageousembodiment. In this example, aligned structure 400 may include compositepart 402 and composite part 404. Composite part 402 may be honeycombsandwich 406, while composite part 404 may be honeycomb sandwich 408.

Honeycomb sandwich 406 may comprise faceplate 410, core 412, andfaceplate 414. Honeycomb sandwich 408 may comprise faceplate 416, core418, and faceplate 420. Faceplate 410, faceplate 414, faceplate 416, andfaceplate 420 may be composite materials in the form of laminates.

Core 412 and core 418 may be honeycomb cores with cells aligned in thedirection of arrow 422. Aligned structure 400 may take various forms.For example, without limitation, aligned structure 400 may be a lightshield, a wall panel, a composite panel, a divider panel, a floor panel,or some other suitable structure. A light shield may be a cover that maybe placed over a light source. The light shield may diffuse light from alight source and/or provide a protective barrier.

With reference now to FIG. 5, a diagram of a cross-sectional viewillustrating layers of materials for joining composite parts is depictedin accordance with an advantageous embodiment. In this example, adhesivelayer 500 may be placed on surface 502 of aligned structure 400.Composite material 504 may be placed onto adhesive layer 500. In asimilar fashion, adhesive layer 506 may be placed on surface 508 ofaligned structure 400. Composite material 510 may be placed ontoadhesive layer 506. Of course, in some advantageous embodiments,adhesive layer 506 and composite material 510 may be unnecessary,depending on the particular implementation.

With reference now to FIG. 6, a diagram of a cross-sectional viewillustrating a completed structure is depicted in accordance with anadvantageous embodiment. In FIG. 6, decorative layer 600 may be appliedto aligned structure 400 after adhesive layer 500, composite material504, adhesive layer 506, and composite material 510 have been cured.

With reference now to FIGS. 7-11, diagrams illustrating joiningcomposite parts to each other are depicted in accordance with anadvantageous embodiment. With reference first to FIG. 7, compositehoneycomb panel 700 and composite honeycomb panel 702 may be alignedwith each other to form aligned structure 704. This alignment may formsplice area 706 with splice line 708.

In this illustration, adhesive layer 710 may take the form of epoxy film712. Epoxy film 712 may be placed onto top side 714 of aligned structure704 in splice area 706 to cover splice line 708.

Turning next to FIG. 8, a diagram illustrating application of anadhesive layer is depicted in accordance with an advantageousembodiment. In this example, epoxy film 800 may be applied to bottomside 802 of aligned structure 704 over splice line 708 in splice area706.

With reference now to FIG. 9, a diagram illustrating an application of acomposite material to a structure is depicted in accordance with anadvantageous embodiment. In this operation, composite material 900 maytake the form of reinforced prepreg 902. Reinforced prepreg 902 may beplaced onto epoxy film 712 on top side 714 of aligned structure 704.

With reference now to FIG. 10, a diagram illustrating application of acomposite material to a structure is depicted in accordance with anadvantageous embodiment. In this illustration, reinforced prepreg 1000may be applied to bottom side 802 of aligned structure 704 over spliceline 708 in splice area 706.

In FIG. 11, a diagram illustrating an aligned structure on a tool isdepicted in accordance with an advantageous embodiment. In thisillustrative example, aligned structure 704 with epoxy film 712, epoxyfilm 800 (not shown), reinforced prepreg 902, and reinforced prepreg1000 (not shown) on splice line 708 in splice area 706 is illustrated onmatch mold die 1100. Match mold die 1100 is an example of oneimplementation of tool 322 in FIG. 3.

Match mold die 1100 includes part 1102 and part 1104. Match mold die1100 may be designed to fit over a portion of aligned structure 704.Match mold die 1100 may cover splice line 708 and splice area 706. Matchmold die 1100 may be heated to apply heat to splice area 706. Match molddie 1100 also may apply pressure to splice area 706.

Turning to FIG. 12, a diagram illustrating a cured aligned structure isdepicted in accordance with an advantageous embodiment. In this example,aligned structure 704 with epoxy film 712 and reinforced prepreg 902 ontop side 714 may have been cured to join composite honeycomb panel 700to composite honeycomb panel 702. With reference to FIG. 13, alignedstructure 704 may be seen with decorative layer 1300 applied to top side714.

The different features and/or operations illustrated in FIGS. 7-13 arenot meant to imply limitations to the manner in which compositecomponents may be joined to each other. For example, in otheradvantageous embodiments, other types of composite components other thanhoneycomb panels may be joined to each other. In yet other advantageousembodiments, layers of materials rather than film adhesives or tape maybe used to apply the different materials to splice area 706.

With reference now to FIG. 14, a flowchart of a process for joiningcomposite parts is depicted in accordance with an advantageousembodiment. The process illustrated in FIG. 14 may be implemented in asplicing environment such as, for example, splicing environment 300 inFIG. 3.

The process may begin by aligning a first composite part with a secondcomposite part to form an aligned structure with a splice line(operation 1400). The process may then place an adhesive layer on thesplice line (operation 1402). This adhesive layer may be, for example,without limitation, an epoxy film. The process may place a compositematerial on the adhesive layer (operation 1404). The composite materialmay be, for example, without limitation, a reinforced prepreg.

The aligned structure with the adhesive layer and the composite materialmay be cured to form a joined structure (operation 1406). A decorativelayer may then be applied to the joined structure to form a completedstructure (operation 1408), with the process terminating thereafter.

The illustration of operations in FIG. 14 is not meant to limit themanner in which different advantageous embodiments may be implemented.In other advantageous embodiments, other operations in addition to, orin place of, the ones illustrated may be used. Further, in someadvantageous embodiments, some operations may be unnecessary. Forexample, the placing of the adhesive layer and the composite materialmay be on both sides of an aligned structure in some advantageousembodiments. In still other advantageous embodiments, the alignedstructure with the adhesive layer and the composite material may beplaced into a die prior to the curing operation.

Thus, the different advantageous embodiments provide a method andapparatus for joining composite parts to each other. The differentadvantageous embodiments may provide a quicker process for joiningcomposite parts. Further, the different advantageous embodiments alsomay provide for better joining or splicing of parts and/or less out oftolerance occurrences as compared to currently used processes.

The description of the different advantageous embodiments has beenpresented for purposes of illustration and description, and it is notintended to be exhaustive or limited to the embodiments in the formdisclosed. Many modifications and variations will be apparent to thoseof ordinary skill in the art. Further, different advantageousembodiments may provide different advantages as compared to otheradvantageous embodiments.

Although the different advantageous embodiments have been described withrespect to aircraft, other advantageous embodiments may be applied toother types of objects. For example, without limitation, otheradvantageous embodiments may be applied to a mobile platform, astationary platform, a land-based structure, an aquatic-based structure,a space-based structure, and/or some other suitable object.

More specifically, the different advantageous embodiments may be appliedto, for example, without limitation, a submarine, a bus, a personnelcarrier, a tank, a train, an automobile, a spacecraft, a space station,a satellite, a surface ship, a power plant, a dam, a manufacturingfacility, a building, and/or some other suitable object.

The embodiment or embodiments selected are chosen and described in orderto best explain the principles of the embodiments, the practicalapplication, and to enable others of ordinary skill in the art tounderstand the disclosure for various embodiments with variousmodifications as are suited to the particular use contemplated.

1. A method for joining composite parts, the method comprising: aligninga first composite part with a second composite part to form an alignedstructure with a splice line; placing an adhesive layer on the spliceline; placing a composite material on the adhesive layer; curing thealigned structure with the adhesive layer and the composite material toform a joined structure.
 2. The method of claim 1, wherein the curingstep comprises: placing the aligned structure with the adhesive layerand the composite material in a die; and heating the aligned structurein the die.
 3. The method of claim 2, wherein the heating stepcomprises: heating the aligned structure in the die with the alignedstructure under pressure.
 4. The method of claim 1 further comprising:applying a decorative layer to the joined structure.
 5. The method ofclaim 1, wherein the aligned structure has as a first side and a secondside, wherein the adhesive layer and the composite material are placedon the first side and the second side, and wherein the adhesive layerand the composite material are placed on the second side.
 6. The methodof claim 2, wherein the die covers the splice line after the adhesivelayer and the composite material have been placed on the splice line. 7.The method of claim 1, wherein the adhesive layer is selected from oneof an adhesive film and an epoxy film.
 8. The method of claim 1, whereinthe aligned structure is an aircraft structure.
 9. The method of claim1, wherein the composite material is a fabric with an impregnated resin.10. The method of claim 9, wherein the fabric is selected from one of afiberglass fabric and a carbon fabric.
 11. The method of claim 9,wherein the resin is selected from one of a phenolic resin and an epoxyresin.
 12. The method of claim 1, wherein the joined structure isselected from one of a composite panel, a wall panel, divider panel, afloor panel, and a light shield.
 13. The method of claim 1, wherein thefirst composite part is a first honeycomb sandwich panel and the secondpart is a second honeycomb sandwich panel.
 14. A method formanufacturing a composite honeycomb sandwich panel, the methodcomprising: aligning a first composite honeycomb sandwich panel with asecond composite honeycomb sandwich panel to form an aligned structurewith a splice line; placing an epoxy film on the splice line, whereinthe epoxy film covers the splice line; placing a fabric with an embeddedresin on the epoxy film, wherein the aligned structure has a first sideand a second side, wherein the epoxy film and the fabric with theembedded resin are placed on the first side and the second side, whereinthe fabric is selected from one of a fiberglass fabric and a carbonfabric, and wherein the resin is selected from one of a phenolic resinand an epoxy resin; placing the aligned structure with the epoxy filmand the fabric with the embedded resin in a die, wherein the die coversthe splice line; heating the aligned structure in the die under pressureto form a joined structure; and applying a decorative layer on thejoined structure, wherein the completed structure is selected from oneof a composite panel, a wall panel, and a light shield.
 15. An apparatuscomprising: a first composite part aligned with a second composite partto form an aligned structure with a splice line; an adhesive layercovering the splice line; a composite material located on the adhesivelayer; a tool capable of holding the aligned structure with the adhesivelayer covering the splice line and the composite material located on theadhesive layer.
 16. The apparatus of claim 15, wherein the adhesivelayer is selected from one of an adhesive film and an epoxy film. 17.The apparatus of claim 17, wherein the aligned structure is an aircraftpart.
 18. The apparatus of claim 15, wherein the composite material is afabric with an impregnated resin.
 19. The apparatus of claim 18, whereinthe fabric is selected from one of a fiberglass fabric and a carbonfabric.
 20. The apparatus of claim 18, wherein the resin is selectedfrom one of a phenolic resin and an epoxy resin.
 21. The apparatus ofclaim 15, wherein the joined structure is selected from one of acomposite panel, a wall panel, divider panel, a floor panel, and a lightshield.
 22. The apparatus of claim 15, wherein the first composite partis a first honeycomb sandwich panel, and the second composite part is asecond honeycomb sandwich panel.
 23. An apparatus for splicing honeycombcomposite parts, the apparatus comprising: a first composite honeycombsandwich panel aligned with a second composite honeycomb sandwich panelto form an aligned structure with a splice line; an epoxy film locatedon the splice line; a fabric with an embedded resin located on the epoxyfilm; a tool capable of holding the aligned structure with the epoxyfilm located on the splice line and the fabric with the embedded resin,wherein the tool is a match mold die capable of receiving the alignedstructure and capable generating pressure on the aligned structure.